The present invention relates generally to gas turbine engines, and, more specifically, to turbine rotor blades therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine which powers the compressor. Additional energy is extracted from the gases in a low pressure turbine which powers a fan in a typical aircraft turbofan gas turbine engine application.
Engine efficiency increases as combustion gas temperature increases, but the gas temperature must be limited for protecting the various components over which the combustion gases flow during operation. For example, the combustion gases are initially confined by the liners of the combustor and channeled between the stator vanes of the turbine nozzle bounded by inner and outer bands. The combustion gases flow between the turbine rotor blades and are bound by radially inner platforms integral therewith and radially outer turbine shrouds surrounding the row of rotor blades.
Each component of the engine is specifically designed with a specific configuration for its specific purpose associated with the hot combustion gases. The hot engine components directly exposed to the hot combustion gases are typically cooled by using a portion of the pressurized air diverted from the compressor which is channeled through corresponding cooling circuits of the components.
The variety of cooling circuits and features thereof is remarkably large due to the associated problems in cooling the variously configured components. Turbine component life is typically limited by local affects, and therefore each component must be specifically designed in toto for protection from the hot combustion gases while maintaining suitable strength of the component for the desired useful life of the component.
Component life is a significant factor in designing modern aircraft turbofan engines which directly affects acquisition and maintenance costs of thereof. Accordingly, state-of-the-art high strength superalloy materials are commonly used in the design of modern aircraft engines, notwithstanding their correspondingly high cost. Superalloy materials, such as nickel or cobalt based superalloys, maintain high strength at high temperature and are desirable in the manufacture of the various hot components of the engine.
In a typical high pressure, first stage turbine rotor blade, the superalloy material thereof is typically enhanced by coating the exposed, external surface of the blade with a thermal barrier coating (TBC). Such coatings are typically ceramic materials which have enhanced thermal insulating performance for protecting the superalloy metallic substrates of the hot components, such as the turbine blade.
The blade includes suitable internal cooling circuits through which the compressor air coolant is channeled for maintaining the operating temperature of the blade below a desired limit for ensuring the intended life for the blade. The blade cooling circuits are myriad in view of the complexity of the airfoil thereof and the corresponding complex temperature distribution of the combustion gases which flow thereover during operation.
Internal cooling circuits typically include dedicated circuits for the leading edge region of the airfoil, the trailing edge region of the airfoil, the mid-chord region of the airfoil, as well as the radially outer tip portion of the airfoil which defines a relatively small clearance or gap with the surrounding turbine shroud. Internal cooling of the airfoil is complemented by external cooling of the airfoil provided by various holes or apertures which extend through the pressure or suction sidewalls, or both, of the airfoil.
The airfoil sidewalls typically include inclined film cooling apertures extending therethrough which discharge the spent cooling air in thin films along the external surface of the airfoil for providing an additional thermal insulating barrier between the airfoil and the hot combustion gases. The variety of film cooling holes themselves is also myriad in view of the complexity of the combustion flowstream surrounding the airfoil. A suitable pressure drop must be provided at each of the film cooling holes to provide a corresponding backflow margin for the holes, as well as discharging the film cooling air without excessive velocity which could lead to undesirable blowoff.
Since the various portions of the airfoil have different operating environments in the combustion gas flow field, they require different cooling configurations. The cooling configurations for the leading edge of the airfoil therefore is not appropriate for the cooling configuration for the trailing edge of the airfoil, and vice versa. Furthermore, the generally concave pressure side of the airfoil operates differently than the generally convex suction side of the airfoil, and correspondingly require different cooling configurations.
And, the radially outer tip of the airfoil typically includes small squealer ribs extending outwardly from the perimeter of the tip which define a small tip cavity above a solid floor of the tip. The combustion gases necessarily leak over the airfoil tip in the clearance provided with the turbine shroud and therefore subject the small squealer ribs to hot combustion gases on both sides thereof. Accordingly, tip cooling requires special configurations, which again are found with myriad differences in conventional applications.
One exemplary gas turbine engine has enjoyed many, many years of successful commercial operation in a marine application. Marine and industrial gas turbine engines are typically derived from their previous turbofan aircraft gas turbine engine parents, and are modified for use in the non-aircraft configurations. These various gas turbine engines nevertheless share common core engines including the compressor, combustor, and high pressure turbine, notwithstanding their different low pressure turbine configuration for providing output power for the fan in the turbofan application or drive shafts in marine and industrial applications.
Although the exemplary marine engine disclosed above has enjoyed many, many thousands of hours of successful commercial use, that long experience has uncovered a form of thermally induced distress in the high pressure, first stage turbine rotor blades nearing the end of their useful lives. In particular, both the blade tip, and the mid-span region of the blade on the suction sidewall just aft of the blade leading edge are showing thermal distress which leads to the degradation of the thermal barrier coating.
Accordingly, it is desired to provide a turbine rotor blade having improved cooling for specifically addressing the newly uncovered local distress in high-time rotor blades.